This invention relates to the field of aerospace systems, and in particular to a general-purpose thermal design that is suitable for a variety of spacecraft missions.
The design of modules for use in a spacecraft conventionally requires a substantial effort in the field of thermal management. At any particular point in time, for example, one of the surfaces of a spacecraft may be facing the sun, and in a typical low-earth orbit receives well over a kilowatt of solar energy per each square meter of surface area facing the sun. At the same time, another surface of the spacecraft may be facing deep-space, at a temperature near zero degrees Kelvin.
U.S. Pat. No. 5,372,183, “THERMAL CONTROL ARRANGEMENTS FOR A GEOSYNCHRONOUS SPACECRAFT”, issued 13 Dec. 1994 to Harold P. Strickberger, and incorporated by reference herein, presents an overview of conventional thermal control systems, and teaches a system wherein spacecraft components are mounted on specified “north” and “south” surfaces, which are defined as the surfaces that face deep space throughout an orbit cycle, and heat pipes are provided to reduce the temperature differential between these “north” and “south” panels. Other, east and west pointing, surfaces are provided with highly reflective surfaces to assure that the absorbed energy is substantially less than the reflected energy on these surfaces.
In the other thermal control systems disclosed in U.S. Pat. No. 5,372,183, information is also available with regard to the expected orientation(s) of the spacecraft throughout the mission. For example, the teachings of U.S. Pat. No. 4,880,050, “THERMAL MANAGEMENT SYSTEM” issued 14 Nov. 1989 to Nakamura et al., and incorporated by reference herein, is described as being well suited for spacecraft that continually rotate, and the teachings of U.S. Pat. No. 3,749,156, “THERMAL CONTROL SYSTEM FOR A SPACECRAFT MODULAR HOUSING” issued 17 Apr. 1972 to Fletcher et al. and incorporated by reference herein, is described as a technique wherein each surface of the spacecraft that is expected to face the sun comprises super-conducting material, to prevent the transfer of heat to other surfaces.
Although mission-specific information regarding the orientation of the spacecraft relative to the sun is generally available during the design of the spacecraft, the need to depend upon such information to design a spacecraft module substantially hinders the design of ‘general-purpose’ modules that can be used on multiple spacecraft, and particularly hinders or precludes the design of modules that are independent of any particular mission.
It is an object of this invention to provide a thermal system for spacecraft modules that allows the modules to be used in a wide range of differing mission-profiles. It is a further object of this invention to provide a method of designing and assembling a spacecraft module that does not require a priori knowledge of the spacecraft module's orientation in an operational environment.
These objects, and others, are achieved by a thermal management system that includes a baseplate and wall system that is precharacterized to provide a given level of thermal performance regardless of an orientation of the spacecraft relative to the sun. The system is characterized at a worst-case hot orientation, and at a worst-case cold orientation. The characterization provides a maximum temperature and a minimum temperature of components mounted on the baseplate as a function of the height of the walls. The height of the walls is selected to provide a suitable temperature range for the components, based on the power dissipation of the components. The system is designed to be symmetric, so that this temperature range is assured regardless of the orientation of the spacecraft.
Throughout the drawings, the same reference numerals indicate similar or corresponding features or functions. The drawings are included for illustrative purposes and are not intended to limit the scope of the invention.